This invention relates to the structure of a liquid-fueled rocket engine, and, more particularly, to the structure of the combustion chamber and its manufacture, and the joining of the propellant injector and the combustion chamber.
A typical liquid-fueled rocket engine includes a cylindrically symmetric combustion chamber, with an injector attached to its injector end and a flared nozzle attached to its nozzle end. A liquid propellant including fuel and an oxidizer flows through injector ports in the injector and into the combustion chamber. The propellant is ignited in the combustion chamber. The hot gas resulting from the combustion expands through the nozzle and drives the rocket engine and the attached rocket structure in the direction opposite to that in which the nozzle is pointed.
The combustion chamber is cylindrically symmetric, but it is known to provide an outward step in the inner wall of the combustion chamber. That is, near the injector the inner surface of the wall of the combustion chamber has a first diameter, but at an axial location a portion of the distance from the injector to the nozzle there is an abrupt outward step in the inner surface of the wall to a second, greater diameter. The presence of this step promotes improved mixing and combustion of the fuel and the oxidizer, leading to improved performance of the rocket engine.
The wall of the combustion chamber is exposed to high temperature combustion gas during service. The wall is preferably made of a refractory material such as rhenium coated with iridium on the inwardly facing surface. The injector plate is much cooler in service, and is typically made of titanium. When the rocket engine is fired, there is a large temperature increase from room temperature and a large temperature gradient between the upper end of the combustion chamber and the adjacent injector, through the region where the two are attached.
It is conventional practice to attach the injector plate to the combustion chamber with a flange-and-bolt system or by welding. The flange-and-bolt system has the disadvantage that the mechanical seal is not well suited to withstand, without leaking, the high temperatures, large temperature change between room temperature and the service temperature, and large temperature gradients during service. Welding of the extremely dissimilar metals is difficult, and in particular it is quite difficult to achieve reliable welds of rhenium to titanium with available welding technology. The high temperatures on the combustion chamber side of the joint and the high temperature gradient through the joint can lead to a premature failure of the joint and a shortening of the life of the engine. Even a small pinhole in the joint can be disastrous, as it results in a back leak of hot combustion gases from the interior of the combustion chamber.
Recent advances in the design of the rocket engine to allow higher-temperature combustion and the use of more powerful propellants have resulted in even greater temperatures and temperature gradients. The existing stepped combustion chamber and combustion chamber/injector structures may be insufficient for operation in this environment. There is therefore a need for an improved approach to the manufacturing and structure of the combustion chamber, and for the attachment of the injector to the combustion chamber. The present invention fulfills this need, and further provides related advantages.